1. Technical Field
The present invention relates to gas turbine engines and, more particularly, to cooling turbine section components and to reducing the potential for a stall or a surge therefor.
2. Background Art
Conventional gas turbine engines are enclosed in an engine case and include a compressor, a combustor, and a turbine. An annular flow path extends axially through the sections of the engine. As is well known in the art, the compressor includes alternating rows of stationary airfoils (vanes) and rotating airfoils (blades) that apply force to compress the incoming working medium. A portion of the compressed working medium enters the combustor where it is mixed with fuel and burned therein. The products of combustion, or: hot gases, then flow through the turbine. The turbine includes alternating rows, of stationary vanes and rotating blades that extend radially across the annular flow path and expand the hot gases to extract force therefrom. A portion of the extracted energy is used to drive the compressor.
Each airfoil includes a low pressure side (suction side) and a high pressure side (pressure side) extending radially from a root to a tip of the airfoil. To optimize efficiency, the annular flow path for the working medium is defined by an outer shroud and an inner shroud. The outer shroud is typically the engine case disposed radially outward of the outer tips of the rotating blades. A tip clearance is defined between the engine case and the tips of the rotating blades.
One of the major goals in gas turbine engine fabrication is to optimize efficiency and performance, without sacrificing engine stability. In order to optimize the efficiency of the compressor and the turbine it is necessary to ensure that work performed on the working medium is not lost. One factor effecting total efficiency is tip leakage losses. Tip leakage occurs when higher pressure air from the pressure side of the rotor blade leaks to the lower pressure suction side of the blade through the tip clearance. Tip leakage reduces efficiency in two ways. First, work is lost when higher pressure gas escapes through the tip clearance without being operated on in the intended manner by the blade, i.e. for compressors the leakage flow is not adequately compressed and for the turbines the leakage is not adequately expanded. Second, leakage flow from the pressure side produces interference with the suction side flow. The interference results from the leakage flow being misoriented with respect to the suction side flow. The difference in the orientation and velocity of the two flows results in a mixing loss as the two flows merge and eventually become uniform. Both types of losses contribute to reduction in efficiency of the gas turbine engine.
Tip leakage also may result in engine instability, such as a stall or surge. If the tip clearance flow is overly strong and sufficiently penetrates into the incoming flow, the direction of the air flow through the compressor will reverse, degrading performance of that stage and potentially causing a surge. Since engine instability is highly undesirable, particularly in aircraft applications, the problem of tip leakage and instability has been investigated for many years. One solution is to reduce the tip clearances and ensure that the engine is operated well below the surge line. Most current solution attempts to reduce tip: clearance involve actively changing the tip clearance by adjusting the diameter of the engine case liner. However, the active control of the tip clearance requires additional hardware that adds complexity and undesirable weight to the engine. Solutions for improving surge line conditions include engine case treatments or bleeding valves, or both, as disclosed in the U.K. Patent Application GB 2158879: entitled xe2x80x9cPreventing Surge in Axial Flow Compressorxe2x80x9d, published Nov. 20, 1985. The U.K. Patent Application discloses use of both case treatments and a bleed valve for selective bleeding of the compressor air. However, the scheme potentially improves engine stability, but sacrifices performance by wasting the compressor air bled through the bleeding valve.
Another factor that effects gas turbine engine performance is the need to cool certain turbine components. The turbine section of the gas turbine engine is subjected to an extremely harsh environment, characterized by very high temperatures and pressures. The components of the turbine must be cooled to prevent these components from burning in a very short period of time. The cooler air is typically bled from the compressor and routed to the turbine. Although the bled cooling air is necessary to cool turbine components, the loss of the cooling air from the compressor is highly undesirable.
Typically, the cooling air bled from the compressor must have pressure high enough to flow downstream to the turbine, but also to be taken from the compressor stage with the lowest pressure usable for cooling purposes such that no additional work is performed on the air, thereby wasting energy and lowering the gas turbine engine efficiency. In order to ensure that no additional work is performed on the extracted cooling air, other than absolutely necessary, the cooling air is diverted from the compressor before the air enters a blade stage and after the air passes through the vane stage. This is done to prevent the following stages of blades from performing additional work on the air and to raise static pressure of the air as it passes through the vane stage. Thus, the need to divert air from the compressor to cool turbine components reduces the overall engine efficiency, but is necessary for the engine performance.
Therefore, it is desirable to improve stability of the gas turbine engine without sacrificing performance thereof.
It is an object of the present invention to improve both performance and stability of the gas turbine engine.
According to the present invention, a gas turbine engine having a compressor, a combustor, and turbine enclosed in an engine case with the compressor having a plurality of alternating rows of rotating blades and stationary vanes includes a substantially circumferential groove formed in the engine case of the compressor substantially adjacent to a row of rotating blades to extract a portion of tip leakage flow from that row of blades and to route the extracted tip leakage flow to the turbine section for cooling turbine components. In the preferred embodiment of the present invention, the groove communicates with a plenum which communicates with the turbine section of the gas turbine engine via channeling pipes formed within the engine case. The extracted tip leakage flow reduces the overall tip leakage flow that would otherwise be available to pass from a pressure side of the blade to a suction side and to mix and interfere with the suction side flow.
The reduction in tip leakage optimizes both, engine efficiency and stability. Efficiency and performance of the gas turbine engine are optimized because first, there is no performance penalty since the bled flow is used for cooling the turbine components and second, tip clearance flow is reduced, thereby reducing interference between the pressure side flow and the suction side flow. Moreover, engine stability is improved since tip leakage ceases to be sufficiently strong to cause a surge. Therefore, the present invention improves not only the efficiency and performance of the gas turbine engine, but also the stability thereof.
The foregoing and other advantages of the present invention become more apparent in light of the following detailed description of the exemplary embodiments thereof, as illustrated in the accompanying drawings.